Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade receive air from the compressor of the turbine engine and pass the air through the blade. Some of the cooling fluids are passed through the root and into the cavity between adjacent turbine blades to cool the platforms of the blades. The cooling fluids may be exhausted through gaps between adjacent blades and may create film cooling. Various seals have been used to limit the flow of cooling fluids between the gap and to limit the influx of hot combustion gases through the gap.
Cooling air and hot gas leakages have a detrimental impact on performance of a gas turbine engine, nitrogen oxides (NOx) emissions and mechanical integrity of components. Air leakage detracts from turbine performance because energy is expended to compress air without incurring any benefit in turbine airfoil cooling and hence in metal temperature reduction. Excessive air leakage from disc cavities may disrupt the flow in the turbine airfoil channels, increase losses and decrease stage efficiency. Also, the more air is extracted from the compressor and dumped into the gas path downstream of the combustor, the higher the primary zone temperature in the combustor has to be for the required engine firing temperature. This results in increased NOx production. Hot gas ingestion into the turbine disc cavities leads to higher disc and blade root temperatures and may result in reduced service lives and failures. In addition, hot gas leakage through the blade root serrations and under the stator assemblies reduces turbine performance. Such losses are exacerbated in engines with increased firing temperatures and pressure ratios. For the same seal clearance, the increase in leakage is directly proportional to the increase in the pressure upstream of the seal. The pressure ratio increase is further driven by gas turbine applications.
Depending on the turbine disc cavity configuration, clearances at the disc rim and flow conditions in the gas path, the disc will pump a specific amount of air flow. If the supplied disc cooling air flow is not able to satisfy this amount, there will be hot gas ingestion from the gas path. As previously set forth, the air and hot gas leakages have a detrimental effect on the turbine performance and the mechanical integrity of the affected parts.